Electrical resistance wear indicator

ABSTRACT

In combination a wear indicator and a component of a gas turbine engine is provided. The wear indicator is secured to a surface of the component of the gas turbine engine. The wear indicator comprising: a first component including: a first plate; a second plate opposite the first plate; a plurality of wires extending from first plate to the second plate, wherein the first plate is electrically connected to the second plate through the plurality of wires; and a potting material configured to partially fill the first component and fill voids between the plurality of wires, such that the plurality of wires are electrically insulated from each other by the potting material.

BACKGROUND

The subject matter disclosed herein generally relates to measurementdevices and, more particularly, to a method and an apparatus fordetecting blade tip clearance for a gas turbine engine.

Gas turbine engines typically include a compressor, a combustor, and aturbine, with an annular flow path extending axially through each.Initially, air flows through the compressor where it is compressed orpressurized. The combustor then mixes and ignites the compressed airwith fuel, generating hot combustion gases. These hot combustion gasesare then directed from the combustor to the turbine where power isextracted from the hot gases by causing blades of the turbine to rotate.

The compressor and turbine sections include multiple rotors and statorsconfigured to enable optimal operation. Gas turbine engines maintain anoptimal clearance (distance) between the tips of the rotors and anoutside diameter of a gas path within the turbine engine, and therebyprovide the conditions necessary to achieve a desired performance.

SUMMARY

According to one embodiment, in combination a wear indicator and acomponent of a gas turbine engine is provided. The wear indicator issecured to a surface of the component of the gas turbine engine. Thewear indicator comprising: a first component including: a first plate; asecond plate opposite the first plate; a plurality of wires extendingfrom first plate to the second plate, wherein the first plate iselectrically connected to the second plate through the plurality ofwires; and a potting material configured to partially fill the firstcomponent and fill voids between the plurality of wires, such that theplurality of wires are electrically insulated from each other by thepotting material.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the first component isconfigured to delaminate when impacted by a blade of the gas turbineengine.

In addition to one or more of the features described above, or as analternative, further embodiments may include a second component of thewear indicator having a blind hole partially enclosing the firstcomponent, wherein the wear indicator is secured to a surface of thecomponent of the gas turbine engine through the second component.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the component is ablade outer air seal.

In addition to one or more of the features described above, or as analternative, further embodiments may include a measurement deviceelectrically connected to the first plate through a first lead line andelectrically connected to the second plate through second lead line,wherein the measurement device is configured to determine the resistanceof at least one of the first plate, the second plate, and the pluralityof wires.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the second componentfurther comprises: a first side that delaminates when impacted by ablade of the gas turbine engine; a second side parallel to the firstside, the second side being secured to the component of the gas turbineengine; and a mid-section interposed between the first side and thesecond side; wherein the mid-section is conical frustum in shape.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the first componenthas a cylindrical shape.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the first plate andthe second plate compose a substantial portion of a curved surface ofthe cylindrical shape, wherein the first plate and the second plate areseparated by a gap in the curved surface.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the first platefurther includes a plurality of first orifices and the second platefurther includes a plurality of second orifices, and wherein each of theplurality of wires extends from a first orifice to a second orifices.

According to another embodiment, in combination a wear indicator and acomponent of a gas turbine engine is provided. The wear indicator issecured to a surface of the component of the gas turbine engine. Thewear indicator comprising: a first component including: a post; and aribbon wire having a first end and a second end opposite the first end,wherein the first end is operably connected to the post, and wherein theribbon wire is wrapped around the post.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the first component isconfigured to delaminate when impacted by a blade of the gas turbineengine.

In addition to one or more of the features described above, or as analternative, further embodiments may include: a second component of thewear indicator having a blind hole partially enclosing the firstcomponent, wherein the wear indicator is secured to a surface of thecomponent of the gas turbine engine through the second component.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the component is ablade outer air seal.

In addition to one or more of the features described above, or as analternative, further embodiments may include a measurement deviceelectrically connected to the first end through a first lead line andelectrically connected to the second end through second lead line,wherein the measurement device is configured to determine the resistanceof the ribbon wire.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the second componentfurther comprises: a first side that delaminates when impacted by ablade of the gas turbine engine; a second side parallel to the firstside, the second side being secured to the component of the gas turbineengine; and a mid-section interposed between the first side and thesecond side; wherein the mid-section is conical frustum in shape.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the post iscylindrical in shape.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the ribbon wire iswrapped around an outer surface of the post.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the ribbon wire iswrapped around an inner surface of the post.

According to another embodiment, a method of detecting blade clearancein a gas turbine engine is provided. The method comprising: attaching awear indicator to an inner surface of a gas turbine engine opposite ablade of the gas turbine engine; measuring a first resistance of thewear indicator; operating the gas turbine engine at a first selectedspeed for a first period of time to remove material from the wearindicator; measuring a second resistance of the wear indicator aftermaterial is removed from the wear indicator; determining a change inresistance between the second resistance and the first resistance; anddetermining an amount of material removed from the wear indicator by theblade in response to the change in resistance.

In addition to one or more of the features described above, or as analternative, further embodiments may include determining a clearancebetween the blade and the inner surface in response to the amount ofmaterial removed from the wear indicator.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, that the followingdescription and drawings are intended to be illustrative and explanatoryin nature and non-limiting.

BRIEF DESCRIPTION

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, like elements are numberedalike:

FIG. 1 is a partial cross-sectional illustration of a gas turbineengine, in accordance with an embodiment of the disclosure;

FIG. 2 is a cross-sectional illustration of the wear indicator, inaccordance with an embodiment of the disclosure;

FIG. 3 is a perspective of a first component of the wear indicator ofFIGS. 2 and 3, in accordance with an embodiment of the disclosure;

FIG. 4a-4g is a flow chart illustrating a method of manufacturing thewear indicator of FIGS. 2-3, in accordance with an embodiment of thedisclosure;

FIG. 5 is a cross-sectional illustration of the wear indicator, inaccordance with an embodiment of the disclosure;

FIG. 6a-6g is a flow chart illustrating a method of manufacturing thewear indicator of FIG. 5, in accordance with an embodiment of thedisclosure; and

FIG. 7 is a flow chart illustrating a method of detecting bladeclearance within a gas turbine engine, in accordance with an embodimentof the disclosure.

The detailed description explains embodiments of the present disclosure,together with advantages and features, by way of example with referenceto the drawings.

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosedapparatus and method are presented herein by way of exemplification andnot limitation with reference to the Figures.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct, while the compressor section 24 drives air along a coreflow path C for compression and communication into the combustor section26 then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. An engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The engine staticstructure 36 further supports bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis Awhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and35,000 ft (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low correctedfan tip speed” as disclosed herein according to one non-limitingembodiment is less than about 1150 ft/second (350.5 m/sec).

Referring now to FIG. 2, which shows a cross-sectional view of a rubbutton or wear indicator 100 installed in a gas turbine engine 20, inaccordance with an embodiment of the disclosure. As seen in FIG. 2, thewear indicator 100 is attached to an inner surface 72 of the gas turbineengine 20 opposite a blade 55 of the gas turbine engine 20. The blade 55rotates along a blade path BP1. In the illustrated embodiment, the wearindicator 100 is attached to a blade outer air seal 62, the outer airseal 62 is the inner surface 72. One or more wear indicators 100 may beaffixed to the inner surface 72 of the gas turbine engine 20 in order tomonitor the clearance between the blade 55 and the inner surface 72, amethod discussed further below in FIG. 7. In the embodiment of FIG. 2,the wear indicator 100 has been installed in the high pressure turbine54 of the gas turbine engine 20. It is understood that the wearindicator 100 may be located in other sections of the gas turbine engine20 having rotating blades 55. An abradable coating 74 may be applied onthe inner surface 72 of the gas turbine engine 20 and the wear indicator100 may be covered by the abradable coating 74 on the inner surface 72.The abradable coating 74 is designed to provide protection for the innersurface 72 against a blade 55 strike. If a blade 55 were to extendtowards the inner surface 72 then the abradable coating 74 shall bestruck first and absorb the impact of the blade 55 to prevent damage tothe inner surface 72. The wear indicator 100 may be attached to theinner surface 72 using an adhesive (not shown) that may or may not needa curing to adhere the wear indicator 100 to the inner surface 72.

As seen in FIG. 2, the wear indicator 100 may comprise a first component130 and a second component 160 configured to partially enclose the firstcomponent 130. The first component includes a plurality of wires 140extending from a first plate 132 to a second plate 134, as seen in FIG.3. The first plate 132 includes a plurality of first orifices 133 andthe second plate 134 includes a plurality of second orifices 135. Theplurality of second orifices 135 are complimentary to the plurality offirst orifices 133. Each wire 140 extends from a first orifice 133 to asecond orifice 135. The plurality of wires 140 extend across a cavity141 between the first plate 132 and the second plate 134. The firstcomponent 130 may be partially filled with a potting material 136. Thepotting material 136 is configured to fill voids 142 between each of theplurality of wires 140, such that the plurality of wires 140 areelectrically insulated from each other by the potting material 136. Thepotting material 136 is non-conductive and capable of withstanding thehigh temperatures of a gas turbine engine. The potting material 136 mayonly partially fill the cavity 141 a portion D1. The first component 130may have a round shape and/or a cylindrical shape, as seen in FIG. 3. Itis understood that the round and/or cylindrical shape of the firstcomponent 130 show in FIG. 3 is not intended to be limiting and thefirst component 130 may have a variety of different shapes. The firstplate 132 and the second plate 134 compose a substantial portion of acurved surface 139 of the cylindrical shape of the first component 130.It is also understood that the curved shape of the plates 132, 134 showin FIG. 3 is not intended to be limiting and the plates 132, 134 mayhave a variety of different shapes. The first plate 132 and the secondplate 134 are separated by the gap D2 in the curved surface 139.Advantageously, the gap D2 keeps the first plate 132 electricallyseparate from the second plate 134.

The first plate 132 may be connected to a first lead line 152 and thesecond plate 134 may be connected to a second lead line 154. The leadlines 152, 154 may be connected to a measurement device 300 configuredto measure the resistance through the plates 132, 134 and plurality ofwires 140. The measurement device 300 may include a processor and amemory. For ease of illustration, the processor and memory are not shownin FIG. 3. The processor can be any type or combination of computerprocessors, such as a microprocessor, microcontroller, digital signalprocessor, application specific integrated circuit, programmable logicdevice, and/or field programmable gate array. The memory is an exampleof a non-transitory computer readable storage medium tangibly embodiedin or operably connected to the path determination system includingexecutable instructions stored therein, for instance, as firmware.

Each of the plates 132, 134 have a known resistance and each of theplurality of wires 140 have a known resistance. Advantageously, sincethe plurality of wires 140 and the plates 132, 134 each have a knownresistance then as the blade 55 cuts into the wear indicator 100removing some of the plurality of wires 140 then the depth that theblade 55 cut into the wear indicator 100 may be determined by measuringthe resistance after the cut and comparing to the original resistanceprior to the cut.

The second component 160 includes a blind hole 162 configured topartially enclose the first component 130. The blind hole 162 includes ablind hole base 163. The first component 130 is inserted into the blindhole 162 such that the second component 160 partially encloses the firstcomponent 130 and the plurality of wires 140 are located proximate thebase 163 of the blind hole 162. In an embodiment, the blind hole 162substantially matches a shape of the first component 130. In anembodiment, the blind hole 162 may be a round shape and/or a cylindricalshape. The first component 130 may be securely attached to the secondcomponent 160 by an epoxy capable of withstanding the high temperaturesof a gas turbine engine 20. The second component 160 may be composed ofa ceramic material capable of withstanding the high temperatures of agas turbine engine 20.

The second component 160 may have a first side 164 and a second side 166parallel to the first side 164. The first side 164 shares a common wall169 with the blind hole base 163. The second component 160 alsocomprises a mid-section 170 interposed between the first side 164 andthe second side 166. In an embodiment, a second diameter D4 of thesecond side 166 may be larger than a first diameter D3 of the first side164. The second side 166 may be affixed to the blade outer air seal 62and/or the abradable coating 74. The second side 166 may be affixed tothe blade outer air seal 62 and/or the abradable coating 74 by an epoxyin a non-limiting example. If the blade 55 strikes the second component160 then a layer of the second component 160 will be removed from thefirst side 164. Thus, the first side 164 delaminates when impacted by ablade 55 of the gas turbine engine 20. Delaminate may be understood tomean the removal of material from the second component 160 in layers. Inan embodiment, the first component 130 is configured to delaminate whenimpacted by a blade 55 of the gas turbine engine 20. The secondcomponent 160 will continue to delaminate in layers until the firstcomponent 130 is exposed to the blade 55 and then the first component130 and the second component 160 will delaminate together. In anembodiment, the mid-section 170 of the second component 160 between thefirst side 164 and the second side 166 may have conical frustum shape,as seen in FIG. 2. Advantageously, since the mid-section 170 is withinthe flow path of the gas turbine engine 20, a conical frustum shape isaerodynamic and may provide reduce disturbance to airflow through thegas turbine engine 20. Also advantageously, a conical frustrum shape mayminimize major material loss when struck by the blade 55. As may beappreciated by one of skill in the art, the second component 160 mayinclude various shapes, sizes and reference dimensions not disclosedherein.

The second component 160 may also include an extrusion 168 projectingout from the second side 166. The extrusion 168 may provide additionalsupport to the first component 130. The extrusion 168 may protrude intothe blade outer air seal 62. A portion of the first component 130 mayprotrude past the blade outer air seal 62, such that the plates 132, 134may be connected to the lead lines 152, 154 without interfering with theblade outer air seal 62.

Referring now to FIGS. 4a-4g with continued reference to FIGS. 1-3.FIGS. 4a-4g shows a flow chart illustrating a method 400 formanufacturing a wear indicator 100 in accordance with an embodiment ofthe present disclosure. Blocks 404, 506, and 508 illustrate thefabrication/assembly of a first component 130 of the wear indicator 100.At block 404, a first plate 132 is formed having a plurality of firstorifices 133. Also at block 404, a second plate 134 is formed having aplurality of second orifices 135. Further at block 404, the second plate134 is oriented in relation to the first plate 132 such that theplurality of second orifices 135 are complimentary to the plurality offirst orifices 133, as seen in FIG. 4. The term complimentary means thatthe orifices 135, 133 are generally in line with each other.

At block 406 a plurality of wires 140 are welded to the first plate 132.Each of the wires 140 is located in a first hole 133 of the first plate132. Also at block 406, each of the plurality of wires 140 extend acrossa cavity 141 between the first plate 132 and the second plate 134.Further at block 406, the plurality of wires 140 are welded to thesecond plate 134. Each of the wires 140 is located in a second hole 135of the second plate 134.

At block 408, a portion D1 of the cavity 141 is filled with a pottingmaterial 136 such that the potting material 136 fills voids 142 betweeneach of the plurality of wires 140. The potting material 136 may enterthe cavity as a liquid flowing in between the wires 140 and then hardento a solid. The plurality of wires 140 are electrically insulated fromeach other by the potting material 136.

At block 410, a second component 160 having a blind hole 162 with ablind hole base 163 is obtained. Also at block 410, the first component130 is inserted into the blind hole 162 such that the second component160 partially encloses the first component 130 and the plurality ofwires 140 are located proximate the base 163 of the blind hole 162.Block 410 may also include forming the second component 160. Asmentioned above, a formed second component 160 may include a first side164 that delaminates when impacted by a blade 55 of the gas turbineengine 20; and a second side 166 parallel to the first side 164. Thefirst side 164 shares a common wall 169 with the blind hole base 163.Block 410 may also include attaching the first component 130 to thesecond component 160. The first component 130 may be attached to thesecond component 160 using an epoxy.

At block 412, the second side 166 is attached to a blade outer air seal62. At block 414, first lead line 152 is electrically connected to thefirst plate 132 and a second lead line 154 is electrically connected tothe second plate 134. At block 416, a portion of the first side 164 ofthe second component 160 is removed after the first component 130 hasbeen inserted into the second component 160. The portion may be removedby grinding the first side 164. Advantageously, a portion of the firstside 164 may be grinded away in order to achieve a desired starting sizefor the wear indicator 100.

While the above description has described the flow process of FIGS.4a-4g in a particular order, it should be appreciated that unlessotherwise specifically required in the attached claims that the orderingof the steps may be varied.

Referring now to FIG. 5, which shows a cross-sectional view of a rubbutton or wear indicator 200 installed in a gas turbine engine 20, inaccordance with an embodiment of the disclosure. As seen in FIG. 5, thewear indicator 200 is attached to an inner surface 72 of the gas turbineengine 20 opposite a blade 55 of the gas turbine engine 20. The blade 55rotates along a blade path BP1. In the illustrated embodiment, the wearindicator 200 is attached to a blade outer air seal 62, the outer airseal 62 is the inner surface 72. One or more wear indicators 200 may beaffixed to the inner surface 72 of the gas turbine engine 20 in order tomonitor the clearance between the blade 55 and the inner surface 72, amethod discussed further below in FIG. 7. In the embodiment of FIG. 5,the wear indicator 200 has been installed in the high pressure turbine54 of the gas turbine engine 20. It is understood that the wearindicator 200 may be located in other sections of the gas turbine engine20 having rotating blades 55. An abradable coating 74 may be applied onthe inner surface 72 of the gas turbine engine 20 and the wear indicator200 may be covered by the abradable coating 74 on the inner surface 72.The abradable coating 74 is designed to provide protection for the innersurface 72 against a blade 55 strike. If a blade 55 were to extendtowards the inner surface 72 then the abradable coating 74 shall bestruck first and absorb the impact of the blade 55 to prevent damage tothe inner surface 72. The wear indicator 200 may be attached to theinner surface 72 using an adhesive (not shown) that may or may not needa curing to adhere the wear indicator 200 to the inner surface 72.

FIG. 5 shows an alternate embodiment of a wear indicator 200 as opposedto the wear indicator that was shown in FIGS. 2-4. As seen in FIG. 5,the wear indicator 200 may comprise a first component 230 and a secondcomponent 260 configured to partially enclose the first component 230.The first component 230 may include a post 250 and a ribbon wire 240. Asseen in FIG. 5, the ribbon wire 240 has a first end 232 and a second end234 opposite the first end. The first end 232 is operably connected tothe post 250 and the ribbon wire 240 is wrapped around the post 250.Once wrapped around the post 250, the ribbon wire 240 forms a coiledresistive element having a known resistance. The ribbon wire 240 mayinclude an insulating material configured to insulate portions of theribbon wire 240 from other portions of the ribbon wire 240 when wrappedin a coiled resistive element, such that the ribbon wire 240 iselectrically insulated from itself. In two non-limiting examples, theinsulating material may be an insulating material on one side of theribbon wire 240 and/or a solidifying insulating liquid that fills thegaps between the ribbon wire 240 when coiled. In an embodiment, the post250 is cylindrical in shape. As seen in FIG. 5, the first end 232 isradially inward from the second end 234 once wrapped, however otherwrapping arrangements may also be used, such as, for example, the post250 may include a larger orifice 254 and the ribbon wire 240 may bewrapped around the inside of the post 250 within the orifice 254, thuswrapping from a radially outward position to a radially inward position.In the illustrated embodiment, the ribbon wire 240 is wrapped around anouter surface 250 a of the post 250. In an alternative embodiment, theribbon wire 240 may be wrapped around the inner surface 250 b of thepost 250. The post 250 may be operably connected to a base 252. The base252 may be oriented at a first angle α1 relative to the post 250. Thefirst angle α1 may be equal to about 90°, thus the base 252 may be aboutperpendicular to the post 250. The base 252 may be circular in shape. Inan embodiment, the post 250 and the base 252 may be composed ofnon-conductive material.

The first end 232 may be connected to a first lead line 152 and thesecond end 234 may be connected to a second lead line 154. The leadlines 152, 154 may be connected to a measurement device 300 configuredto measure the resistance through the ribbon wire 240. The measurementdevice 300 may include a processor and a memory. For ease ofillustration, the processor and memory are not shown in FIG. 5. Theprocessor can be any type or combination of computer processors, such asa microprocessor, microcontroller, digital signal processor, applicationspecific integrated circuit, programmable logic device, and/or fieldprogrammable gate array. The memory is an example of a non-transitorycomputer readable storage medium tangibly embodied in or operablyconnected to the path determination system including executableinstructions stored therein, for instance, as firmware.

The ribbon wire 240 has a known resistance, as mentioned above.Advantageously, since the ribbon wire 240 has a known resistance, as theblade 55 cuts into the wear indicator 200 removing a portion of theribbon wire 240 then the depth that the blade 55 cut into the wearindicator 200 may be determined by measuring the resistance after thecut and comparing to the original resistance prior to the cut.

The second component 260 includes a blind hole 262 configured topartially enclose the first component 230. The blind hole 262 includes ablind hole base 263. The first component 230 is inserted into the blindhole 262 such that the second component 260 partially encloses the firstcomponent 230 and the ribbon wire 240 is located proximate the base 263of the blind hole 262. In an embodiment, the blind hole 262substantially matches a shape of the first component 230. In anembodiment, the blind hole 262 may be a round shape and/or a cylindricalshape. The first component 230 may be securely attached to the secondcomponent 260 by an epoxy capable of withstanding the high temperaturesof a gas turbine engine 20. The second component 260 may be composed ofa ceramic material capable of withstanding the high temperatures of agas turbine engine 20.

The second component 260 may have a first side 264 and a second side 266parallel to the first side 264. The first side 264 shares a common wall269 with the blind hole base 263. The second component 260 alsocomprises a mid-section 270 interposed between the first side 264 andthe second side 266. In an embodiment, a second diameter D7 of thesecond side 266 may be larger than a first diameter D6 of the first side264. The second side 266 may be affixed to the blade outer air seal 62and/or the abradable coating 74. The second side 266 may be affixed tothe blade outer air seal 62 and/or the abradable coating 74 by an epoxyin a non-limiting example. If the blade 55 strikes the second component260 then a layer of the second component 260 will be removed from thefirst side 264. Thus, the first side 264 delaminates when impacted by ablade 55 of the gas turbine engine 20. Delaminate may be understood tomean the removal of material from the second component 260 in layers. Inan embodiment, the first component 230 is configured to delaminate whenimpacted by a blade 55 of the gas turbine engine 20. The secondcomponent 260 will continue to delaminate in layers until the firstcomponent 230 is exposed to the blade 55 and then the first component230 and the second component 260 will delaminate together. In anembodiment, the mid-section 270 of the second component 260 between thefirst side 264 and the second side 266 may have conical frustum shape,as seen in FIG. 5. Advantageously, since the mid-section 270 is withinthe flow path of the gas turbine engine 20, a conical frustum shape isaerodynamic and may provide reduce disturbance to airflow through thegas turbine engine 20. Also advantageously, a conical shape may minimizemajor material loss when struck by the blade 55. As may be appreciatedby one of skill in the art, the second component 260 may include variousshapes, sizes and reference dimensions not disclosed herein.

The second component 260 may also include an extrusion 268 projectingout from the second side 266. The extrusion 268 may provide additionalsupport to the first component 230. The extrusion 268 may protrude intothe blade outer air seal 62.

Referring now to FIGS. 6a-6g with continued reference to FIG. 5. FIGS.6a-6g shows a flow chart illustrating a method 600 for manufacturing awear indicator 200 in accordance with an embodiment of the presentdisclosure. Blocks 604 and 608 illustrate the fabrication/assembly of afirst component 230 of the wear indicator 200. At block 604, a post 250is formed. The post 250 may be operably connected to a base 252. Thepost 250 may also include an orifice 254 and a slot 256. The orifice 254may run parallel along the length L1 of the post 250, as seen in FIG. 6a. The slot 256 may cut the post 250 in half along the length L1 of thepost 250, as seen in FIG. 6a . At block 606, a first end 232 of theribbon wire 240 is inserted into the slot 256 of the post 250 and afirst lead line 152 is electrically connected to the first end 232within the orifice 254, as seen in FIG. 6b . At block 608, the ribbonwire 240 is wrapped around the post 250. Once wrapped around the post250, the ribbon wire 240 forms a coiled resistive element having a knownresistance. At block 610 a second lead line 154 is electricallyconnected to the second end 234 of the ribbon wire 240.

At block 612, a second component 260 having a blind hole 262 with ablind hole base 263 is obtained. Also at block 612, the first component230 is inserted into the blind hole 262 such that the second component260 partially encloses the first component 230 and ribbon wire 240 islocated proximate the base 263 of the blind hole 262. Block 612 may alsoinclude forming the second component 260. As mentioned above, a formedsecond component 260 may include a first side 264 that delaminates whenimpacted by a blade 55 of the gas turbine engine 20; and a second side266 parallel to the first side 264. The first side 264 shares a commonwall 269 with the blind hole base 263. Block 612 may also includeattaching the first component 230 to the second component 260. The firstcomponent 230 may be attached to the second component 260 using anepoxy.

At block 614, the second side 266 is attached to a blade outer air seal62. At block 616, a portion of the first side 264 of the secondcomponent 260 is removed after the first component 230 has been insertedinto the second component 260. The portion may be removed by grindingthe first side 264. Advantageously, a portion of the first side 264 maybe grinded away in order to achieve a desired starting size for the wearindicator 200.

While the above description has described the flow process of FIGS.6a-6g in a particular order, it should be appreciated that unlessotherwise specifically required in the attached claims that the orderingof the steps may be varied.

Referring now to FIG. 7 with continued reference to FIGS. 1-6. FIG. 7 isa flow chart illustrating a method 700 for detecting blade clearance ina gas turbine engine 20, in accordance with an embodiment. At block 706,the wear indicator 100, 200 is attached to an inner surface 72 of a gasturbine engine 20 opposite a blade 55 of the gas turbine engine 20. Atblock 710, a first resistance of the wear indicator 100, 200 ismeasured. At block 712, the gas turbine engine 20 is operated at a firstselected speed for a first period of time to remove material from thewear indicator 100, 200. At block 714, a second resistance of the wearindicator 100, 200 is measured after material is removed from the wearindicator 100, 200. At block 716, a change in resistance between thesecond resistance and the first resistance is determined. At block 718,an amount of material removed from the wear indicator 100, 200 by theblade 55 is determined in response to the change in resistance. Method600 may also include: determining a clearance between the blade 55 andthe inner surface 72 in response to the amount of material removed fromthe wear indicator 100, 200.

While the above description has described the flow process of FIG. 7 ina particular order, it should be appreciated that unless otherwisespecifically required in the attached claims that the ordering of thesteps may be varied.

Technical effects of embodiments of the present disclosure include usinga wear indicator to determine blade tip clearance through detecting achange in electrical resistance.

The term “about” is intended to include the degree of error associatedwith measurement of the particular quantity based upon the equipmentavailable at the time of filing the application. For example, “about”can include a range of ±8% or 5%, or 2% of a given value.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a”, “an” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof.

While the present disclosure has been described with reference to anexemplary embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. In combination a wear indicator and a componentof a gas turbine engine, wherein the wear indicator is secured to asurface of the component of the gas turbine engine, the wear indicatorcomprising: a first component including: a first plate; a second plateopposite the first plate; a plurality of wires extending from firstplate to the second plate, wherein the first plate is electricallyconnected to the second plate through the plurality of wires; and apotting material configured to partially fill the first component andfill voids between the plurality of wires, such that the plurality ofwires are electrically insulated from each other by the pottingmaterial.
 2. The combination of claim 1, wherein the first component isconfigured to delaminate when impacted by a blade of the gas turbineengine.
 3. The combination of claim 1, further comprising: a secondcomponent of the wear indicator having a blind hole partially enclosingthe first component, wherein the wear indicator is secured to a surfaceof the component of the gas turbine engine through the second component.4. The combination of claim 1, wherein the component is a blade outerair seal.
 5. The combination of claim 1, further comprising ameasurement device electrically connected to the first plate through afirst lead line and electrically connected to the second plate throughsecond lead line, wherein the measurement device is configured todetermine the resistance of at least one of the first plate, the secondplate, and the plurality of wires.
 6. The combination of claim 3,wherein the second component further comprises: a first side thatdelaminates when impacted by a blade of the gas turbine engine; a secondside parallel to the first side, the second side being secured to thecomponent of the gas turbine engine; and a mid-section interposedbetween the first side and the second side; wherein the mid-section isconical frustum in shape.
 7. The combination of claim 1, wherein thefirst component has a cylindrical shape.
 8. The combination of claim 7,wherein the first plate and the second plate compose a substantialportion of a curved surface of the cylindrical shape, wherein the firstplate and the second plate are separated by a gap in the curved surface.9. The combination of claim 1, wherein the first plate further includesa plurality of first orifices and the second plate further includes aplurality of second orifices, and wherein each of the plurality of wiresextends from a first orifice to a second orifice.
 10. In combination awear indicator and a component of a gas turbine engine, wherein the wearindicator is secured to a surface of the component of the gas turbineengine, the wear indicator comprising: a first component including: apost; and a ribbon wire having a first end and a second end opposite thefirst end, wherein the first end is operably connected to the post, andwherein the ribbon wire is wrapped around the post.
 11. The combinationof claim 10, wherein the first component is configured to delaminatewhen impacted by a blade of the gas turbine engine.
 12. The combinationof claim 10, further comprising: a second component of the wearindicator having a blind hole partially enclosing the first component,wherein the wear indicator is secured to a surface of the component ofthe gas turbine engine through the second component.
 13. The combinationof claim 10, wherein the component is a blade outer air seal.
 14. Thecombination of claim 10, further comprising a measurement deviceelectrically connected to the first end of the ribbon wire through afirst lead line and electrically connected to the second end of theribbon wire through a second lead line, wherein the measurement deviceis configured to determine the resistance of the ribbon wire.
 15. Thecombination of claim 12, wherein the second component further comprises:a first side that delaminates when impacted by a blade of the gasturbine engine; a second side parallel to the first side, the secondside being secured to the component of the gas turbine engine; and amid-section interposed between the first side and the second side;wherein the mid-section is conical frustum in shape.
 16. The combinationof claim 10, wherein: the post is cylindrical in shape.
 17. Thecombination of claim 10, wherein: the ribbon wire is wrapped around anouter surface of the post.
 18. The combination of claim 10, wherein: theribbon wire is wrapped around an inner surface of the post.
 19. A methodof detecting blade clearance in a gas turbine engine, the methodcomprising: attaching a wear indicator to an inner surface of a gasturbine engine opposite a blade of the gas turbine engine; measuring afirst resistance of the wear indicator; operating the gas turbine engineat a first selected speed for a first period of time to remove materialfrom the wear indicator; measuring a second resistance of the wearindicator after material is removed from the wear indicator; determininga change in resistance between the second resistance and the firstresistance; and determining an amount of material removed from the wearindicator by the blade in response to the change in resistance.
 20. Themethod of claim 19, further comprising: determining a clearance betweenthe blade and the inner surface in response to the amount of materialremoved from the wear indicator.